This invention relates generally to gas turbine engines, and more specifically to methods and apparatus for assembling gas turbine engine components.
Accurate fabrication of engine components may be a significant factor in engine performance and engine efficiency. Specifically, when the component is a gas turbine engine blade, the fabrication of the blade may affect the overall performance and efficiency of the gas turbine engine. At least some known gas turbine engines include high and low pressure compressors, a combustor, and at least one turbine. The compressors compress air which is mixed with fuel and channeled to the combustor. The fuel/air mixture is then ignited to generate hot combustion gases, which are channeled to the turbine. At least some known turbines include a rotor assembly that includes at least one row of circumferentially-spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side coupled to a suction side at a leading edge and a trailing edge. Each airfoil extends radially outward from a rotor blade platform. At least some known rotor blades include a dovetail that extends radially inward from a shank coupled to the platform. The dovetail is used to mount the rotor blade within the rotor assembly to a rotor disk or spool. In at least some known gas turbine engines, a small gap may be defined between a lower surface of the dovetail and a lower surface of the rotor disk.
During operation, a pressure differential created between the rotor blade pressure side and the rotor blade suction side may result in an undesirable leakage flow between the upstream and downstream portions of the rotor. One such possible leakage path may be defined through the gap defined between the dovetail and the lower surface of the rotor disk groove in which the rotor blades are carried. If such leakage paths are not efficiently sealed, the leakage flow may have an adverse effect both on engine efficiency and engine performance.